This invention relates generally to gas turbine engines, and in particular, to a cooled flow path surface region.
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000xc2x0 F., considerably higher than the melting temperatures of the metal parts of the engine which are in contact with these gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying a cooling air to the outer surfaces of the metal parts through various methods. The metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are the metal parts forming combustors and parts located aft of the combustor including turbine blades and vanes and exhaust nozzles.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet gases is normally limited by the materials used to fabricate the components downstream of the combustors such as the vanes and the blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to 2100xc2x0-2200xc2x0 F.
The metal temperatures can be maintained below melting levels with current cooling techniques by using a combination of improved cooling designs and thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air from the compressor internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades have intricate serpentine passageways within the structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from a hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engines, a process known as film cooling.
In another approach, a thermal barrier coating (TBC) is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component.
TBCs include well-known ceramic materials, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not adhere well directly to the superalloys used as substrate materials. Therefore, an additional metallic layer called a bond coat is placed between the substrate and the TBC. The bond coat may be made of a nickel containing overlay alloy, such as a NiCrAlY or a NiCoCrAlY, or other composition more resistant to environmental damage than the substrate, or alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide whose surface oxidizes to a protective aluminum oxide scale that provides improved adherence to the ceramic top coatings. The bond coat and overlying TBC are frequently referred to as a thermal barrier coating system.
While superalloys coated with such thermal barrier coating systems do provide substantially improved performance over uncoated materials, there remains room for improvement. Improved environmental resistance to destructive oxidation and hot corrosion is desirable. In some instances, the alloying elements of the bond coat interdiffuse with the substrate alloy and consume walls of the turbine airfoils, i.e., reduces load carrying capability, thereby limiting blade life. This interdiffusion can also reduce environmental resistance of the coating.
Even with the use of advanced cooling designs and thermal barrier coatings, it is also desirable to decrease the requirement for cooling air, because reducing the demand for cooling air also contributes to improving overall engine operating efficiency. One way to achieve such a reduction is to improve the cooling of the metal parts immediately adjacent to their outer surfaces.
U.S. Pat. No. 5,195,243 to Junod is directed to a method of making a coated porous metal panel. A plurality of metal panels are fastened in such a manner so as to be separated by a plurality of integral, raised pedestals each having a flat bonding surface. An internal chamber is thereby defined, the chamber receiving a flow of cooling air from inlet pores on one metal panel and discharging the cooling air through discharge pores on a second metal panel. The raised pedestals create a tortuous flow path for the cooling air as it passes from the inlet pores through the internal chamber to the discharge pores. A thermally resistant coating is applied to the panel containing the discharge pores, followed by placement of a shield. The placement of the inlet pores, discharge pores and extraction passages (located directly behind the discharge pores) is critical to minimize deposit of surplus coating material in the internal chamber and in the discharge and inlet pores.
U.S. Pat. No. 5,034,284 to Bornstein et al. discloses a porous strain isolation layer placed between the substrate and the bond coat. The porous layer is formed by spraying a mixture of alloy and polymer particles and subsequently heating to eliminate the polymer. The pores are in a random pattern and do not create channels.
Thus, there is an ongoing need for an improved thermal barrier coating system, wherein the environmental resistance and long-term stability of the thermal barrier coating system is improved so that higher engine efficiencies can be obtained. In previous designs, the bond coat temperature limit is critical to the TBC""s life and has had an upper limit of about 2100xc2x0 F. Once the bond coat exceeds this temperature, the coating system will quickly deteriorate, due to high temperature mechanical deformation and accelerated oxidation as well as a more rapid interdiffusion of elements with the substrate alloy. The coating system can separate from the substrate exposing the underlying superalloy component to damage from the hot gasses.
What is needed are improved designs that will allow turbine engine components to run at higher operating temperatures, thus improving engine performance without additional cooling air. The present invention fulfills this need, and further provides related advantages.
The present invention provides active convection cooling through micro channels within or adjacent to the bond coat layer applied to a turbine engine component, for example, a blade or vane. When placed adjacent to or within a porous TBC, the micro channels additionally provide transpiration cooling through the porous TBC. The micro channels communicate directly with at least one cooling circuit contained within the blade or vane from which they receive cooling air, thereby providing direct and efficient cooling for the bond coat layer. Because the substrate includes an actively cooled flow path surface region that can reduce the cooling requirement for the substrate, the engine can run at a higher firing temperature without the need for additional cooling air, achieving a better, more efficient engine performance.
In one form, the present invention provides for an actively cooled flow path surface region for a gas turbine hot section component comprising at least one channel in the component substrate material having a first and second end. The first end terminates in an exit orifice located at a pre-selected position on a first surface of the substrate material and connected to a plenum, the second end communicating with a cooling fluid supply. At least one micro channel communicates with at least one exit orifice. The micro channel is located at an interface between the first surface of the substrate material and a first surface of a second material, such as a bond coat, and is substantially parallel to the surface of the substrate and substantially perpendicular to the channel in the substrate. Optionally, a ceramic thermal barrier coating (TBC) is layered to the second, opposed surface of the second material.
In another form, the present invention sets forth a method for actively cooling the flow path surface region of an engine component used in a gas turbine engine comprising the steps of masking the surface of the substrate with a masking material, the masking material forming a pattern on the substrate surface overlying at least one cooling fluid supply contained within the component, coating the masked substrate surface with a bond coat; removing the masking material, leaving hollow micro channels in the pattern occupied by the masking material before its removal for the transport of a cooling fluid and passing a cooling fluid through the micro channels.
In still another form, the present invention sets forth a method for actively cooling the flow path surface region of an engine component used in a gas turbine engine comprising the steps of coating a substrate with a bond coat, masking the exposed surface of the bond coat with a masking material to form a pattern on the exposed bond coat surface in communication with at least one cooling fluid supply contained within the engine component, applying a TBC to the masked bond coat, removing the masking material leaving hollow micro channels in the pattern occupied by the masking material before its removal for the transport of a cooling fluid, and optionally passing a cooling fluid through the micro channels.
In still another form, the present invention delineates a method for actively cooling the flow path surface region of an engine component used in a gas turbine engine comprising the steps of coating a substrate with a first bond coat layer, applying a masking material to the exposed surface of the bond coat to form a predetermined pattern on the exposed bond coat surface in communication with at least one cooling fluid supply contained within the engine component, applying a second bond coat layer over the masking material, applying a TBC to the second bond coat layer, removing the masking material, leaving hollow micro channels in the pattern occupied by the masking material before its removal for the transport of a cooling fluid, and optionally passing a cooling fluid through the cooling channels.
In still another form, the present invention delineates a method for actively cooling the flow path surface region of an engine component used in a gas turbine engine comprising the steps of coating a substrate with a bond coat layer, applying a first TBC layer to the bond coat, applying a masking material to the exposed surface of the first TBC layer to form a predetermined pattern in communication with at least one cooling fluid supply contained within the engine component, applying a second, porous TBC layer over the masking material, removing the masking material, leaving hollow micro channels for the transport of a cooling fluid in the pattern occupied by the masking material before its removal, and optionally passing a cooling fluid through the micro channels.
In still another form, the present invention sets forth a method for actively cooling the flow path surface region of an engine component used in a gas turbine engine comprising the steps of grooving a surface of a substrate such that at least one micro groove communicates with at least one cooling fluid supply contained within the engine component, placing a masking material within the micro grooves, placing a bond coat on the grooved substrate, applying a TBC layer over the bond coat, removing the masking material leaving hollow micro channels for the transport of a cooling fluid, and, optionally passing a cooling fluid through the micro channels.
The present invention further comprises the cooled flow path surface region formed by the foregoing methods and the turbine component with the patterned micro channels substantially parallel to the surface of the substrate for cooling the component.
An advantage of the present invention is the flow path surface region of the coated gas turbine component is actively cooled. By removing heat from this region, the integrity of the bond coat can be maintained at higher engine operating temperatures.
In one embodiment, the active convection cooling through the micro channels occurs within or adjacent to the bond coat layer, providing direct and efficient cooling for the bond coat layer. Since the substrate is covered with the bond coat layer, the cooling requirement for the substrate will also be reduced.
Another advantage of the present invention is that the actively cooled bond coat layer will allow engine components to run at higher operating temperatures to achieve a better engine performance.